|Publication number||US7699269 B2|
|Application number||US 12/193,213|
|Publication date||Apr 20, 2010|
|Filing date||Aug 18, 2008|
|Priority date||Mar 11, 2004|
|Also published as||US7424989, US20050224659, US20090200430|
|Publication number||12193213, 193213, US 7699269 B2, US 7699269B2, US-B2-7699269, US7699269 B2, US7699269B2|
|Inventors||Dale M. Pitt|
|Original Assignee||The Boeing Company|
|Export Citation||BiBTeX, EndNote, RefMan|
|Patent Citations (16), Referenced by (2), Classifications (8), Legal Events (3)|
|External Links: USPTO, USPTO Assignment, Espacenet|
This application is a divisional of U.S. patent application Ser. No. 10/798,687 filed on Mar. 11, 2004. The entire disclosure of the above application is incorporated herein by reference.
This disclosure relates generally to flight control systems and, more particularly, to flight control systems for suppressing aerodynamically induced vibrations.
High performance fighter aircraft such as the F-18 and F-22 (available from the Boeing Company of Chicago, Ill.) often experience high frequency aerodynamically induced vibrations of their wings, stabilators, and vertical tails. These vibrations are cause by buffeting aerodynamic forces and are transmitted into, and through, the aircraft structure. If uncompensated for, the associated stresses may lead to premature cracking of the structure. The resulting repairs of these cracks are both expensive and time consuming. In the alternative, aerodynamic modifications to reduce the causative turbulence impose performance constraints on the aircraft while structural modifications to reduce the resulting fatigue stress incur weight and cost penalties.
Additionally, the buffeting of the vehicle transmits noise into, and causes noise within, the aircraft structure. In turn, the structure transmits the noise to the cockpit wherein noise control measures, with additional penalties must be undertaken. Nor are these problems isolated to high performance aircraft. Automobiles, missiles, and launch and reentry vehicles (for example) also receive buffeting from aerodynamic forces.
Various attempts have been made to use the existing flight control actuators to compensate for these aero-vibrations. However, the flight control system typically commands the actuator at rates of about 30 cycles per second or less. Since the vibrations occur at frequencies significantly higher than the commands, such attempts have failed.
Moreover, attempts to modify the flight control system to react quickly enough to respond to the vibration are impractical for a variety of reasons. For instance, such modifications require an order of magnitude increase in the flight computer speed. Thus, these solutions necessitate an upgrade of the computer. Additionally, modifying the flight control laws to accommodate the additional functionality necessitate the recertification of the flight control system. These recertifications are expensive, time consuming, and (as such) highly undesirable.
Furthermore, as composite materials replace aluminum, and other conventional, structural members (e.g. on the Boeing 7E7 aircraft) vibration control may assume an increasing importance in the design, operation, and maintenance of aircraft. Thus a need exists to reduce or eliminate aerodynamically induced vibrations.
It is in view of the above problems that the present disclosure was developed. The disclosure provides apparatus and methods for reducing vibrations of mobile platform structures.
In one embodiment, the disclosure provides a self-contained actuation device that reduces the aerodynamic buffet loads. Accordingly, construction and operation of mobile platforms (e.g. aircraft) in accordance with the principles of the present disclosure results in more cost efficient platforms that possess better performance and longer service lives.
Another embodiment employs the existing flight control actuators to reduce the aero-buffeting without requiring modification of the flight control system. The present embodiment also includes a vibration sensor placed on an aircraft structure (e.g. a wing) to sense the aero-vibration. The sensor is connected to a controller that is coupled to the actuator. The controller inverts the vibration signal and adds it to the actuator command from the flight control system. Then the controller sends the combined signal (the command with the inverted vibration signal superimposed thereon) to the actuator, thereby driving the actuator out of phase with the vibration. By thus canceling the vibration the current embodiment reduces cyclical loads and improves the fatigue life of the structure.
In yet another embodiment, a method of reducing aerodynamically induced vibrations is provided. The method includes sensing the vibration and inverting a signal representative of the same. The inverted vibration signal is superimposed on a flight control system command for an actuator to drive the actuator out of phase with the vibration.
Further features and advantages of the present disclosure, as well as the structure and operation of various embodiments of the present disclosure, are described in detail below with reference to the accompanying drawings.
The accompanying drawings, which are incorporated in and form a part of the specification, illustrate the embodiments of the present disclosure and together with the description, serve to explain the principles of the disclosure. In the drawings:
Referring to the accompanying drawings in which like reference numbers indicate like elements,
During flight, the flight computer and the pilot continually sense the flight related conditions (pitch, climb rate, speed, and the like). Depending on the conditions, either the computer, the pilot, or both issue commands (e.g. electro-magnetic signals) to reposition the ailerons 14 and other control surfaces. These commands cause the actuator 20 to either extend or retract to move the aileron 14. As noted previously, the wing 12 rushing through the air causes turbulence in the air that causes the wing 12 to vibrate. Additionally, the movement of the control surface aileron 14 tends to change the airflow, thus introducing another source of turbulence, buffeting, and vibration.
Referring now to
Additionally, the actuator typically includes a conventional power cable 30 for receiving power from the aircraft 10 power subsystem. A second power cable 32 may branch from the first power cable 30 to supply the circuit 26 power. Importantly, branching the second power cable 32 from cable 30 saves cable weight by obviating the need for a dedicated cable run from the aircraft 10 power subsystem to power the circuit 26. Another cable 30′ carries the flight control system command to the circuit 26, while in previous systems the cable 30′ was connected directly to the actuator instead of the circuit.
Moreover, coupling the vibration sensor 28 to the actuator 20 instead of the wing 12 (
Furthermore, coupling the circuit 26 to the actuator 20 eliminates the need for cables between the circuit and the actuator to carry the command signal from the circuit and to carry the position signal from the actuator. Likewise, because the flight control system already includes a cable 30′ to the actuator (to carry the actuator command) the present disclosure requires no modifications to the flight control system. Accordingly, the present disclosure further reduces integration costs and delays.
With reference now to
To complete the feedback loop, the summing element 102 superimposes the inverted vibration signal on the command. The resulting signal (the command with the inverted vibration signal superimposed on it) is fed to the actuator 104 to drive the control surface out of phase with the vibration. Accordingly, the control surface 106 causes a disturbance that cancels the vibration of the wing 108. Those skilled in the art will recognize that the commands typically operate up to approximately 30 Hz while the vibration occurs at substantially higher frequencies. Similarly, the commands are typically signals having amplitudes well in excess of the amplitude of the vibration signal (or can be so tailored with appropriate choice of system gains). Thus the inverted vibration signal appears as a ripple superimposed upon the command signal after the two are added.
Since many flight control systems are designed to sense the actual position of the actuator 104, a feedback signal may also be provided by the present disclosure. In particular the vibration sensor 110 and the position sensor 116 (of the actuator 104) may communicate with the filter 114 that subtracts the vibration signal from the position signal as sensed by the position sensor. Accordingly, the position signal fed back to the flight control system does not reflect the slight difference between the commanded position and the actual position as influenced by the vibration canceling circuit 126. Thus, by filtering the vibration signal from the actuator position signal, the present disclosure ensures that the flight control system operates properly (e.g. does not raise a false alarm to indicate that the actuator is out of position).
Note should also be made that
Now with reference to
Additionally, a flight control actuator 228 is shown external to the circuit 226. Schematically
Additionally, a housing 234 of the circuit 226 is rigidly coupled to the actuator so that the vibration transducer 212 accurately senses the vibration of the actuator 228. In turn, the actuator 228 is rigidly mounted to the structure for which vibration reduction is desired (e.g. by the attachment means 22 shown by
Thus, in operation, the summing amplifier 202 sums the flight control command 210 and the vibration signal 220 as inverted by the inverting amplifier 204. The summing amplifier 202 outputs the command 214, with the inverted vibration signal superimposed thereon, to the actuator. As noted previously, the actuator 228 causes an aerodynamic disturbance that cancels the vibration of the structure. Subsequently, if another transient disturbance causes the vibration to return, the vibration cancellation loop (as just described) acts to cancel the new vibration. In the alternative, a non-inverted vibration signal could be subtracted from the command to the same general effect of canceling the vibration.
The circuit 226 also includes a position feedback subsystem that includes the difference amplifier 206. The amplifier 206 accepts the position signal 216 from the actuator position sensor and subtracts the position signal 216 from the vibration signal supplied by the sensor 212. Accordingly, the amplifier 206 filters the vibration signal from the position signal 216 and communicates the resulting signal 218 to the flight control system.
With reference now to
With the power on, commands for the actuator may then be received while the vibration is sensed as at 312 and 314. The two signals are added, subtracted, or superimposed to obtain a command with an inverted vibration signal superimposed upon it at operation 315. The combined signal, from operation 315, is used to drive the actuator out of phase with the vibration. See operation 316. Meanwhile, the position of the actuator may be sensed in 318. Accordingly, the vibration signal may be filtered from the position feedback signal and the result forwarded the flight control system. See for instance operation 320. In this manner, the method may repeat for as long as vibration canceling is desired as indicated at decision 322.
In view of the foregoing, it will be seen that the several advantages of the disclosure are achieved and attained. Notably, the embodiments described herein require no added cabling, or cabling modifications outside of the envelope of the actuator. Moreover, the present disclosure requires no modification to the flight control system, or even the flight computer. Thus, the present disclosure provides vibration elimination with a weight and cost savings over previous attempts to reduce aerodynamically induced vibrations.
The embodiments were chosen and described in order to best explain the principles of the disclosure and its practical application to thereby enable others skilled in the art to best utilize the disclosure in various embodiments and with various modifications as are suited to the particular use contemplated. For example, many industrial machines include structures that are moved by an actuator commanded by a control system. If the machine includes a tool (e.g. a drill) on the movable structure, the tool may induce undesired vibrations in the structure. Thus, the apparatus and methods discussed herein may be adapted to sense the tool induced vibration of the structure and cancel the same without requiring modification of the machine control system. Accordingly, machines with increased precision and accuracy are also provided by the present disclosure.
As various modifications could be made in the constructions and methods herein described and illustrated without departing from the scope of the disclosure, it is intended that all matter contained in the foregoing description or shown in the accompanying drawings shall be interpreted as illustrative rather than limiting. Thus, the breadth and scope of the present disclosure should not be limited by any of the above-described exemplary embodiments, but should be defined only in accordance with the following claims appended hereto and their equivalents.
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|U.S. Classification||244/174, 244/99.13|
|International Classification||B64C13/00, B64D1/12, B64C13/16|
|Cooperative Classification||B64C13/16, Y02T50/44|
|Aug 19, 2008||AS||Assignment|
Owner name: THE BOEING COMPANY, ILLINOIS
Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:PITT, DALE M.;REEL/FRAME:021410/0510
Effective date: 20040305
Owner name: THE BOEING COMPANY,ILLINOIS
Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:PITT, DALE M.;REEL/FRAME:021410/0510
Effective date: 20040305
|Oct 21, 2013||FPAY||Fee payment|
Year of fee payment: 4
|Oct 20, 2017||MAFP|
Free format text: PAYMENT OF MAINTENANCE FEE, 8TH YEAR, LARGE ENTITY (ORIGINAL EVENT CODE: M1552)
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